Geared turbomachine fan and compressor rotation

ABSTRACT

An exemplary gas turbine engine includes a fan section including a fan rotor and at least one fan blade. A fan pressure ratio across the at least one fan blade is less than 1.45, noninclusive of the pressure across any fan exit guide vane system. The engine further includes a low-pressure compressor having a low-pressure compressor rotor that rotates together with the fan rotor at a common speed in operation, and a geared architecture that drives the low-pressure compressor rotor and the fan rotor. The geared architecture has a gear reduction ratio of greater than 2.5. The engine further includes a high-pressure compressor having a pressure ratio greater than 20, a low-pressure turbine having a pressure ratio greater than 5, and a bypass ratio greater than 10.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. patent application Ser. No.13/356,940, which was filed on 24 Jan. 2012 and is incorporated hereinby reference.

BACKGROUND

This disclosure relates to a geared turbomachine having a compressorrotor and a fan rotor that rotate together.

Turbomachines, such as gas turbine engines, typically include a fansection, a turbine section, a compressor section, and a combustorsection. Turbomachines may employ a geared architecture connecting thefan section and the turbine section. The compressor section typicallyincludes at least a high-pressure compressor and a low-pressurecompressor. The compressors include rotors that rotate separately from arotor of fan. To maximize performance of such turbomachines, variousrecent engine architectures have been proposed in which the fan rotatesin a first direction and at a first speed as compared to a low-pressurecompressor which rotates in the opposite direction and at a higherspeed. These recent engine architectures can also be improved.

SUMMARY

A high-bypass ratio geared turbomachine according to an exemplary aspectof the present disclosure comprises a compressor section of ahigh-bypass ratio geared turbomachine, the compressor section providingat least a low-pressure compressor and a high-pressure compressor,wherein a rotor of the low-pressure compressor rotates together with arotor of a fan.

In a further non-limiting embodiment of any of the foregoing high-bypassratio geared turbomachine embodiments, the rotor of the low-pressurecompressor and the rotor of the fan may rotate at the same speed and inthe same direction.

In a further non-limiting embodiment of any of the foregoing high-bypassratio geared turbomachine embodiments, the high-bypass ratio gearedturbomachine may have a fan bypass ratio greater than about 8.

In a further non-limiting embodiment of any of the foregoing high-bypassratio geared turbomachine embodiments, the high-bypass ratio gearedturbomachine may have an overall compression ratio greater than about40.

In a further non-limiting embodiment of any of the foregoing high-bypassratio geared turbomachine embodiments, the high-pressure compressor mayhave a pressure ratio greater than about 20.

In a further non-limiting embodiment of any of the foregoing high-bypassratio geared turbomachine embodiments, the fan may include a shaft thatis rotatably supported by a plurality of tapered bearings.

In a further non-limiting embodiment of any of the foregoing high-bypassratio geared turbomachine embodiments, the high-bypass ratio gearedturbomachine may include a turbine shaft that rotates a gearedarchitecture to rotate the rotor of the low-pressure compressor and therotor of the fan.

In a further non-limiting embodiment of any of the foregoing high-bypassratio geared turbomachine embodiments, at least one thrust bearing mayrotatably support the turbine shaft, and the at least one thrust bearingmay be located axially between the geared architecture and a turbinesecured to the turbine shaft.

In a further non-limiting embodiment of any of the foregoing high-bypassratio geared turbomachine embodiments, the at least one thrust bearingmay be a bi-directional tapered bearing.

In a further non-limiting embodiment of any of the foregoing high-bypassratio geared turbomachine embodiments, the shaft may be a low-pressureturbine shaft.

In a further non-limiting embodiment of any of the foregoing high-bypassratio geared turbomachine embodiments, the geared architecture may be aplanetary geared architecture.

A high-bypass ratio turbomachine according to another exemplary aspectof the present disclosure comprises a fan rotor that rotates togetherwith an compressor rotor at a first speed in a turbomachine having ahigh-bypass ratio, wherein the fan rotor and the compressor rotor aredriven by a shaft that rotates at a second speed different than thefirst speed.

In a further non-limiting embodiment of any of the foregoing high-bypassratio geared turbomachine embodiments, the shaft may rotate a gearedarchitecture to rotate the fan rotor and the compressor rotor.

In a further non-limiting embodiment of any of the foregoing high-bypassratio geared turbomachine embodiments, the compressor rotor may beaxially forward of a fan frame extending radially across a fan bypassflow path.

In a further non-limiting embodiment of any of the foregoing high-bypassratio geared turbomachine embodiments, the fan rotor and the compressorrotor may rotate at the same speed.

In a further non-limiting embodiment of any of the foregoing high-bypassratio geared turbomachine embodiments, the high-bypass ratio gearedturbomachine may include a high-pressure turbine, a combustor section,and a low-pressure turbine arranged axially sequentially within thegeared turbomachine.

In a further non-limiting embodiment of any of the foregoing high-bypassratio geared turbomachine embodiments, exclusively axial compressors mayprovide compression in the geared turbomachine.

In a further non-limiting embodiment of any of the foregoing high-bypassratio geared turbomachine embodiments, the compressor may be alow-pressure compressor.

A method of operating a high-bypass ratio turbomachine according to anexemplary aspect of the present disclosure comprises rotating a gearedarchitecture with a first shaft, rotating a second shaft with the gearedarchitecture, and rotating a fan rotor and a compressor rotor with thesecond shaft.

In a further non-limiting embodiment of any of the foregoing methods ofoperating a high-bypass ratio geared turbomachine, the turbomachine mayhave a fan bypass ratio greater than 8.

DESCRIPTION OF THE FIGURES

The various features and advantages of the disclosed examples willbecome apparent to those skilled in the art from the detaileddescription. The figures that accompany the detailed description can bebriefly described as follows:

FIG. 1 shows a highly schematic view of a portion of an exampleturbomachine.

FIG. 2 shows a schematic view of another example turbomachine.

DETAILED DESCRIPTION

Referring to FIG. 1, an example geared turbomachine 10 includes a firstshaft 11 that provides a rotating input to a geared architecture 12.Rotating the geared architecture 12 rotates a second shaft 13. Theexample geared architecture 12 has a gear ratio that causes the secondshaft 13 to rotate at a slower speed than the first shaft 11.

A compressor rotor 14 and a fan rotor 15 are coupled to the second shaft13. Rotating the second shaft 13 rotates the rotors 14 and 15 at thesame rotational speed and in the same direction. In this example, thecompressor rotor 14 forms a portion of an axial compressor.

FIG. 2 schematically illustrates another example turbomachine, which isa gas turbine engine 20 in this example. The gas turbine engine 20 is atwo-spool turbofan gas turbine engine that generally includes a fansection 22, a compressor section 24, a combustion section 26, and aturbine section 28. Other examples may include an augmentor section (notshown) among other systems or features.

In the example engine 20, the fan section 22 drives air along a bypassflowpath while the compressor section 24 drives air along a coreflowpath. Compressed air from the compressor section 24 communicatesthrough the combustion section 26. The products of combustion expandthrough the turbine section 28.

The example engine 20 generally includes a low-speed spool 30 and ahigh-speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36.

Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with turbofans. Thatis, the teachings may be applied to other types of turbomachines andturbine engines including three-spool architectures.

The low-speed spool 30 and the high-speed spool 32 are rotatablysupported by several bearing systems 38 a-38 d. It should be understoodthat various bearing systems 38 a-38 d at various locations mayalternatively, or additionally, be provided.

The low-speed spool 30 generally includes an inner shaft 40 thatinterconnects a geared architecture 48 and a low-pressure turbine rotor46. The inner shaft 40 is a turbine shaft in this example as the innershaft 40 is connected directly to the low-pressure turbine rotor 46.Rotating the geared architecture 48 rotatably drives a fan rotor 42 anda low-pressure compressor rotor 44 at a lower speed than the low-speedspool 30.

The high-speed spool 32 includes an outer shaft 50 that interconnects ahigh-pressure compressor rotor 52 and high-pressure turbine rotor 54.

In this example, the low-pressure compressor rotor 44 and thehigh-pressure compressor rotor 52 are both rotors of axial compressors,and there are no other types of compressors within the compressorsection 24 of the engine 20.

The combustion section 26 includes a circumferentially distributed arrayof combustors 56 generally arranged axially between the high-pressurecompressor rotor 52 and the high-pressure turbine rotor 54.

A mid-turbine frame 58 of the engine static structure 36 is generallyarranged axially between the high-pressure turbine rotor 54 and thelow-pressure turbine rotor 46. The mid-turbine frame 58 supports thebearing systems 38 c and 38 d in the turbine section 28. The mid-turbineframe 58 includes airfoils 60 within the path of the core airflow.

The inner shaft 40 and the outer shaft 50 are concentric and rotate viathe bearing systems 38 b-38 d about the engine central longitudinal axisA, which is collinear with the longitudinal axes of the inner shaft 40and the outer shaft 50.

In the example engine 20, the core airflow is compressed by thecompressor section 24, mixed and burned with fuel in the combustors 56,then expanded within the turbine section 28. The high-pressure turbinerotor 54 and the low-pressure turbine rotor 46 rotatably drive therespective high-speed spool 32 and low-speed spool 30 in response to theexpansion.

In some non-limiting examples, the engine 20 is a high-bypass gearedaircraft engine. In a further example, the engine 20 has a fan bypassratio that is greater than about six (6:1). In a still further example,the engine 20 has a fan bypass ratio that is greater than about eight(8:1). The overall compression ratio of such the example engine 20 isgreater than 40 (40:1) in some examples, and the pressure ratio of thehigh-pressure compressor is greater than 20 (20:1).

The geared architecture 48 of the example engine 20 includes anepicyclic gear train, such as a planetary geared architecture or othergeared architecture. The example epicyclic gear train has a gearreduction ratio of greater than about 2.3 (2.3:1).

The low-pressure turbine pressure ratio is pressure measured prior toinlet of low-pressure turbine as related to the pressure at the outletof the low-pressure turbine (and prior to an exhausting from the engine20). In one non-limiting embodiment, the bypass ratio of the engine 20is greater than about ten (10:1), the fan diameter is significantlylarger than that of the low-pressure compressor, and the low-pressureturbine has a pressure ratio that is greater than about 5 (5:1). Thegeared architecture 48 of this embodiment is an epicyclic gear trainwith a gear reduction ratio of greater than about 2.5 (2.5:1). Examplesof the geared architecture 48 include star architectures and planetaryarchitectures. It should be understood, however, that the aboveparameters are only exemplary of one embodiment of a geared architectureengine and that the present disclosure is applicable to other gasturbine engines including direct drive turbofans.

In some embodiments of the example engine 20, a significant amount ofthrust is provided by the bypass flow B due to the high bypass ratio.The fan section 22 of the engine 20 is designed for a particular flightcondition—typically cruise at about 0.8 Mach and about 35,000 feet. Thisflight condition, with the engine 20 at its best fuel consumption, isalso known as bucket cruise Thrust Specific Fuel Consumption (TSFC).TSFC is an industry standard parameter of fuel consumption per unit ofthrust.

Fan Pressure Ratio is the pressure ratio across a blade of the fansection 22 (without the use of a Fan Exit Guide Vane system). The lowFan Pressure Ratio according to one non-limiting embodiment of theexample engine 20 is less than 1.45 (1.45:1).

Low Corrected Fan Tip Speed is the actual fan tip speed divided by anindustry standard temperature correction of “T”/518.7^(0.5). Trepresents the ambient temperature in degrees Rankine. The Low CorrectedFan Tip Speed according to one non-limiting embodiment of the exampleengine 20 is less than about 1150 fps (351 m/s).

In the example engine 20, the fan rotor 42 and the low-pressurecompressor rotor 44 are directly connected to a shaft 62. One axial endof the shaft 62 is directly connected to a carrier gear 64 of the gearedarchitecture 48. The fan rotor 42 and the low-pressure compressor rotor44 rotate at the same speed and in the same direction with the shaft 62when the shaft 62 is driven by the carrier gear 64 of the gearedarchitecture 48. The shaft 62 is rotated together with the carrier gear64 when the geared architecture 48 is rotatably driven by the innershaft 40 of the low-speed spool 30. The shaft 62 is considered a fanshaft in this example, because the fan rotor 42 is connected to theshaft 62.

Rotating the fan rotor 42 and the low-pressure compressor rotor 44 withthe shaft 62 facilitates positioning the low-pressure compressor of thecompressor section 24 relatively axially close to the fan section 22.The low-pressure compressor rotor 44 (and thus the low-pressurecompressor) is axially forward of a fan frame 68 in this example. Thefan frame 68 extends radially across a fan bypass passage of the engine20. The fan frame 68 supports an outer duct 70 of the engine 20 relativeto an engine core.

Bearings 38 a rotatably support the shaft 62. The bearings 38 a aretapered in this example. Tapered bearings mounted as shown in FIG. 2will react to the fan 42 thrust loads as well as any radial or momentloads applied to shaft 62 which come from fan 42. In another example,bearings 38 a can be a ball and roller bearing combination. Thiscombination will also react any thrust, radial or moment loads from thefan 42 to the shaft 62. One skilled in the art and having the benefit ofthis disclosure may arrive at other bearing configurations that supportreaction loads applied to shaft 62

Other bearings 38 b rotatably support the low-speed spool 30 near thegeared architecture 48. The bearings 38 b are thrust bearings in thisexample. In one specific example, the bearings 38 b are bi-directionaltapered thrust bearings. In another specific example, the bearings 38 bare ball thrust bearings.

Notably, the example bearings 38 b are located axially between thegeared architecture 48 and the low-pressure turbine rotor 46, and arepositioned axially closer to the geared architecture 48 than thelow-pressure turbine rotor 46.

Positioning the bearings 38 b in this area has some performanceadvantages in the unlikely event that the inner shaft 40 fractures.After such a fracture of the inner shaft 40, axially displacing thelow-pressure turbine rotor 46 relative to other portions of the engine20 is often desired. The axial displacement after a fracture will causethe low-pressure turbine rotor 46 to desirably clash.

Fractures of the inner shaft 40 that are axially forward of the bearings38 b may not result in clash because the bearings 38 b (which are thrustbearings) hold the axial position of the fractured portion. Positioningthe bearings 38 b axially near the geared architecture 48 increases theaxial locations aft the bearings 38 b, and thus the potential fracturelocations of the inner shaft 40 that will result in clash. The bearing38 c and 38 d, in this example, would permit axial displacement after afracture.

In some examples, the torsional strength of the inner shaft 40 is lessthan the torsional strength of the other drive shaft within the engine20 (including the geared architecture 48). Thus, in the event of, forexample, an overload of the fan rotor 42, the inner shaft 40 will failbefore other areas of the engine 20.

The preceding description is exemplary rather than limiting in nature.Variations and modifications to the disclosed examples may becomeapparent to those skilled in the art that do not necessarily depart fromthe essence of this disclosure. Thus, the scope of legal protectiongiven to this disclosure can only be determined by studying thefollowing claims.

We claim:
 1. A gas turbine engine, comprising: a fan section including afan rotor and at least one fan blade, with a fan pressure ratio acrossthe at least one fan blade of less than 1.45, noninclusive of thepressure across any fan exit guide vane system; a low-pressurecompressor having a low-pressure compressor rotor that rotates togetherwith the fan rotor at a common speed in operation; a geared architecturethat drives the low-pressure compressor rotor and the fan rotor, thegeared architecture having a gear reduction ratio of greater than 2.5; ahigh-pressure compressor having a pressure ratio greater than 20; alow-pressure turbine having a pressure ratio greater than 5; and abypass ratio greater than
 10. 2. The gas turbine engine of claim 1,wherein the geared architecture is a planetary geared architecture. 3.The gas turbine engine of claim 2, further comprising a low correctedfan tip speed less than 1150 ft/second, wherein the low corrected fantip speed is an actual fan tip speed at a temperature divided by(T/518.7)^(0.5), where T represents the temperature in degrees Rankine.4. The gas turbine engine of claim 3, further comprising a two-stagehigh-pressure turbine.
 5. The gas turbine engine of claim 4, furthercomprising an engine overall compression ratio greater than
 40. 6. Thegas turbine engine of claim 5, further comprising a fan bypass passageand a fan frame extending radially across the fan bypass passage,wherein the low-pressure compressor is axially forward of the fan frame.7. The gas turbine engine of claim 5, further comprising a low-speedspool that interconnects the geared architecture and the low-pressureturbine, and at least one bearing rotatably supporting the low-speedspool and positioned axially closer to the geared architecture betweenthe geared architecture and the low-pressure turbine.
 8. The gas turbineengine of claim 7, wherein the at least one bearing includes at leasttwo bearings that are bi-directional tapered thrust bearings.
 9. The gasturbine engine of claim 5, wherein the low-pressure turbine is athree-stage low-pressure turbine.
 10. The gas turbine engine of claim 9,further comprising a fan bypass passage and a fan frame extendingradially across the fan bypass passage, wherein the low-pressurecompressor is axially forward of the fan frame.
 11. The gas turbineengine of claim 5, wherein the high-pressure compressor is a nine-stagehigh-pressure compressor.
 12. The gas turbine engine of claim 4, furthercomprising a low-speed spool that interconnects the geared architectureand the low-pressure turbine, and at least one bearing rotatablysupporting the low-speed spool and positioned axially closer to thegeared architecture between the geared architecture and the low-pressureturbine.
 13. The gas turbine engine of claim 12, wherein thelow-pressure turbine rotates an inner shaft, and the high-pressurecompressor is rotated by an outer shaft, wherein a torsional strength ofthe inner shaft is less than a torsional strength of the outer shaft.14. The gas turbine engine of claim 1, further comprising a two-stagehigh-pressure turbine.
 15. The gas turbine engine of claim 14, whereinthe geared architecture is positioned between the low-pressurecompressor and the high-pressure compressor.
 16. The gas turbine engineof claim 15, wherein the fan rotor and the low-pressure compressor rotorare directly connected to a fan shaft, and an axial end of the fan shaftis directly connected to a carrier gear of the geared architecture. 17.The gas turbine engine of claim 16, wherein the low-pressure turbineincludes a low-pressure turbine rotor, and the low-pressure turbinedrives the geared architecture through an inner shaft connected directlyto the low-pressure turbine rotor.
 18. The gas turbine engine of claim17, wherein the fan shaft is rotatably supported by a plurality oftapered bearings.
 19. The gas turbine engine of claim 15, wherein thelow-pressure compressor includes a plurality of stages.
 20. The gasturbine engine of claim 19, wherein each of the at least one fan bladeis forward of the low-pressure compressor with respect to an enginecentral longitudinal axis.